AERODYNAMICS OF AN AIRFOIL WITH PLAIN FLAP IN PRESENCE OF MOMENTUM INJECTION

Transcription

1 AERODYNAMICS OF AN AIRFOIL WITH PLAIN FLAP IN PRESENCE OF MOMENTUM INJECTION BENJAMIN I. TRIPLETT B.A.Sc, University of British Columbia, Canada, 2000 A THESIS SUBMITTED IN PARTIAL FULFILMENT OF THE REQUIREMENTS FOR THE DEGREE OF MASTER OF APPLIED SCIENCE in The Faculty of Graduate Studies Department of Mechanical Engineering We accept this thesis as conforming to the required standard THE UNIVERSITY OF BRITISH COLUMBIA June 2002 Benjamin I. Triplett, 2002

2 In presenting this thesis in partial fulfillment of the requirements for an advanced degree at the University of British Columbia, I agree that the Library shall make it freely available for reference and study. I further agree that the permission for extensive copying of this thesis for scholarly purposes may be granted by the head of my department or by his or her representatives. It is understood that copying or publication of this thesis for financial gain shall not be allowed without my written permission. The University of British Columbia Department of Mechanical Engineering 2324 Main Mall Vancouver, B.C., Canada V6T 1Z4

3 ABSTRACT The concept of Moving Surface Boundary-layer Control (MSBC), as applied to a NASA LS(1)-04T7 airfoil with plain flap, is investigated through a planned wind tunnel test-program at a subcritical Reynolds number of The airfoil carries two rotating cylinders for momentum injection. One is located at the leading edge of the wing and the other a the leading edge of the flap. For conciseness, they are referred to as wing cylinder and flap cylinder, respectively. High speed rotating cylinders controlled the key momentum injection parameters, U W AJ and Uf/U. Here U w and Uf are the surface velocities of the wing and flap cylinders, respectively, and U is the free stream wind speed. Experiments are conducted to characterize the performance of both the two-dimensional (i.e. infinite aspect ratio, 2-D), and of the three-dimensional (i.e. finite aspect ratio, 3-D) cases. The 2-D model airfoil with leading edge momentum injection demonstrated the effectiveness of the concept. The maximum lift of the airfoil was increased by 177% with a delay in stall from 10 to 50. In general, the airfoil performance improves with an increase in U w /U. The momentum injection at the flap enhanced the effect of flap deflection. With the flap deflected to 45, the flap cylinder rotation further shifted the lift curve to the left, and increased the maximum lift coefficient from 1.65 to On the other hand, there is no apparent change in the stall angle of attack. Furthermore, the flap cylinder rotation increased the magnitude of the suction pressure peak and decreased the suction pressure at the trailing edge, suggesting a narrowed wake. For the 2-D model with a flap deflection of 45, the combined wing and flap cylinder rotations showed significant improvement on the airfoil performance. The maximum lift of the airfoil in this configuration was 3.6, a 300% increase over the basic airfoil. This is 213% higher ii

4 lift than that for the airfoil with flap deflection but without momentum injection. The leading edge cylinder delays separation of the boundary-layer, allowing the airfoil to operate at high angles of incidence without stall. The effective wing-camber, introduced by flap deflection, is further increased through any combination of the momentum injection at the wing and the flap. The three-dimensional study indicated that the relative effectiveness of momentum injection at the wing and flap, is the same, i.e. it is independent of the effect of aspect ratio. As well, the effect of aspect ratio is essentially the same with and without the MSBC. These facts should help aircraft designers in implementation of the MSBC in prototype designs. The fundamental information presented in the thesis should serve as a reference and prove useful in further study aimed at aerodynamic design of a wing. iii

5 TABLE OF CONTENTS ABSTRACT TABLE OF CONTENTS LIST OF FIGURES LIST OF SYMBOLS ACKNOWLEDGEMENT DEDICATION ii iv vi ix xi xii 1 INTRODUCTION Preliminary Remarks A Brief Review of the Relevant Literature Scope of the Present Investigation 14 2 MODEL AND TEST-PROGRAM Introductory Remarks Two-Dimensional Study Approach to 3-D Study Pressure and Force Measurements 21 3 TWO-DIMENSIONAL STUDY Reference Airfoil MSBC Applied only at the Wing Rotating Cylinder at the Flap Momentum Injection at the Wing and the Flap 43 4 THREE-DIMENSIONAL STUDY Base NASA LS(1)-0417 Airfoil 50 iv

6 4.2 The Airfoil with Wing Cylinder The Airfoil with Flap Deflection and Flap Cylinder Rotation Momentum Injection at the Wing and the Flap 63 5 CONCLUDING REMARKS Research Contributions Conclusions Recommendations for Future Investigations 69 REFERENCES 71 APPENDIX A 74 V

7 LIST OF FIGURES Figure 1-1 Figure 1-2 Figure 1-3 Figure 1-4 Figure 1-5 Figure 1-6 Figure 1-7 Figure 1-8 Figure 2-1 Figure 2-2 Examples of several devices leading to increased lift through delay in stall. Indicated values are arbitrary [1]. 2 "Ship of Zero Resistance" as conceived by Prandtl. Two counter-rotating cylinders in a uniform flow display essentially potential flow character implying small resistance [13]. 4 Flettner applied the Magnus Effect to ship propulsion in 1924 when he fitted two large vertical rotating cylinders on the deck of the "Buchau" [14]. 6 The practical application of a moving wall for boundary-layer control was demonstrated by Favre im 1938 [15]. Using an airfoil with the upper surface partly formed by a belt moving over two rollers, he was able to delay separation until the angle of attack reached 55, where the maximum lift coefficient of 3.5 was realized. 7 The North American Rockwell OV-1 OA aircraft in flight demonstrating the successful application of the rotating cylinder as a high-lift device. 9 Various rotating cylinder configurations studied with the two-dimensional Joukowski airfoil model by Modi et al. [31 ]. 11 Application of the moving surface boundary-layer control to bluff bodies such as a flat plate at large angles of attack, rectangular prisms, and tractor-trailer truck configurations [32,33]. 12 Representative flow visualization pictures taken by Modi et al. showing, rather dramatically, successful control of the boundary-layer separation through momentum injection [32]. 13 Schematic diagrams showing model configurations used during the testprogram: (a) NASA LS(1)-0417 airfoil, camber, 17% thick at 1/ 3 chord; (b) leading edge of the wing formed by a rotating cylinder; (c) the airfoil with a plain flap, leading edge of the flap replaced with a rotating cylinder; (d) a rotating cylinder at the leading edge of the wing and the flap. 18 Photograph of the 2-D model arrangement in the Parkinson Wind Tunnel. 19 Figure 2-3 Schematic diagram of the Parkinson Wind Tunnel. 20 Figure 2-4 Schematic diagram of the Boundary-layer Wind Tunnel. 22 VI

8 Figure 2-5 Photograph of the 3-D model arrangement in the BLWT. The flap is deflected to Figure 2-6 Arrangement for pressure measurements on the wing rotating cylinder. 25 Figure 2-7 Experimental setup for the acquisition of pressure data. 26 Figure 2-8 Figure 3-1 Figure 3-2 Figure 3-3 Photograph of the model showing the motor drive, rotating cylinders, pressure tubes, and a Variac controller. 28 Schematic diagram of the model used to obtain reference information showing: (a) base airfoil; (b) modifications leading to equivalent airfoil. 31 Pressure distribution for the basic airfoil, NASA LS(1)-0417, showing stall around Pressure distribution for the airfoil in presence of the leading edge momentum injection. 34 Figure 3-4 Pressure plots for the airfoil, with the momentum injection of U w /U = 4, at several angles of attack. 35 Figure 3-5 Comparison of pressure integrated and directly measured lift and drag data for the base airfoil. 36 Figure 3-6 C L vs. a for the airfoil with momentum injection over the range of U w / U = Figure 3-7 Figure 3-8 Figure 3-9 Figure 3-10 Figure 3-11 Figure 3-12 Variation of the drag coefficient with angle of attack in presence of momentum injection. 40 Close-up of the low angle of attack range showing small values of C D attained in presence of momentum injection. 41 Pressure plots showing the effect of flap deflection and momentum injection. 42 Effect of flap deflection and cylinder rotation on the variation of C L with a. 44 Effect of wing cylinder rotation on the pressure distribution for the airfoil with Uf/U = 3 and 5 f = Effect of momentum injection at the flap on the pressure distribution for the airfoil with U w /U = 4 and 8f = 45. The pressure distribution for U w /U = 4 and 8f = 0 is included for reference. 46 vii

9 Figure 3-13 Figure 3-14 Figure 4-1 Figure 4-2 Variation of C L with a for the airfoil with both rotating cylinders and flap deflection. 47 Typical drag data for the airfoil in the presence of both rotating cylinders and flap deflection. 49 Variation of the pressure distribution with angle of attack for the base airfoil. 51 Comparison of pressure distributions for 2-D and 3-D base airfoils with a = Figure 4-3 Comparison between 2-D and 3-D lift variations for the base airfoil. 54 Figure 4-4 Pressure distribution for the 3-D airfoil with momentum injection. 55 Figure 4-5 Figure 4-5 Figure D and 3-D pressure plots, with momentum injection of U w /U = 4, as affected by the angle of attack: (a) a = D and 3-D pressure plots, with momentum injection of U w /U = 4, as affected by the angle of attack: (b) a = D and 3-D pressure plots, with momentum injection of U w /U = 4, as affected by the angle of attack: (c) a = Figure 4-6 Variation of lift with angle of attack for the finite aspect ratio wing. 60 Figure 4-7 Figure 4-8 Figure 4-9 Figure 4-10 Figure A-1 Figure A-2 Drag variation with angle of attack for airfoil with the wing cylinder rotation. Reduction in drag due to delay in stall in presence of the MSBC is apparent. 61 Drag variation with angle of attack for the airfoil with momentum injection. 62 Comparison of the lift variation for 2-D and 3-D cases as affected by the flap cylinder rotation. The flap deflection is held fixed at 8f = Comparison of lifting performance for 2-D and 3-D arrangements of airfoil with two rotating elements and flap deflection. 65 Variation of lift with angle of attack for several cases of MSBC and flap deflection 75 Variation of drag with angle of attack for several cases of MSBC and flap deflection 76 viii

10 LIST OF SYMBOLS 2- D two-dimensional 3- D three-dimensional A wing planform area, c x b AR wing aspect ratio, b 2 / A b wing span BLWT Boundary-Layer Wind Tunnel c airfoil chord C D drag coefficient, D / (1/2) p U 2 A C L lift coefficient, L / (1/2) p U 2 A C-L.max maximum lift coefficient Cp mean pressure coefficient, (P - P^) / (1/2) p U 2 D drag force L lift force LE leading edge LS(1)-0417 NASA low speed, high performance airfoil designation MSBC Moving Surface Boundary-layer Control P static pressure at the surface of the model Poo free stream pressure Re Reynolds number, p U c / p, ix

11 U free stream velocity Uc rotating cylinder surface velocity Uf surface velocity of the rotating cylinder at leading edge of the flap U w surface velocity of the rotating cylinder at leading edge of the wing Uf/U momentum injection parameter for the flap U w /U momentum injection parameter for the wing x distance along airfoil chord x / c nondimensional length along airfoil chord a airfoil angle of attack a stall angle of attack at stall ^ a stall change in stall angle of attack ^L.max change in maximum lift coefficient 5f flap deflection angle p air density p. air viscosity x

12 ACKNOWLEDGEMENT I would like to thank my supervisor, Professor V.J. Modi, for his guidance and support, which has been responsible for so much of my graduate education. The model used in the research was fabricated in the Mechanical Engineering machine shop. The assistance of Mr. Niel Jackson, Mr. Dave Camp, Mr. Doug Yuen, Mr. Len Drakes, Mr. Glen Jolly, and Mr. Gordon Wright, in the design and construction of the model, is gratefully acknowledged. I would like to express my sincere appreciation to my colleagues in the Aerodynamics and Controls Laboratory, Yang Cao, Ken Wong, Jian Zhang and Dr. Ayhan Akinturk, for their friendship and help. Finally, many thanks are due to my friend in the Naval Architecture Laboratory, Peter Ostafichuk, for his assistance with instrumentation and software. xi

13 Dedicated to Kimberley xii

14 1. INTRODUCTION 1.1 Preliminary Remarks Aerodynamicists have perennially sought approaches to increase lifting capacity of an aircraft wing. Besides raising the payload, the higher lift can be used to improve the airplane's maneuvering performance. The lifting force for a given airplane speed, wing area and angle of attack, depends on several factors. These include two-and three-dimensional performance of the wing indicated by maximum lift coefficient, stall angle of attack, control surface distribution, aspect ratio and others. Effects of several leading edge devices on the lifting characteristics of a wing are indicated in Figure 1-1 [1]. Here, increase in the lift is associated with a higher stall angle through the delay in the boundary-layer separation. It is apparent that control of the boundary-layer separation is one of the major parameters affecting the pressure distribution on a wing and its lifting characteristics. A variety of approaches such as suction, blowing, vortex generators, turbulence promotors, and others, have been studied at length and used in practice with varying degrees of success depending on the nature of the application [2]. However, the Moving Surface Boundary-layer Control (MSBC) has received relatively less attention. This is indeed surprising, as the Associate Committee on Aerodynamics, appointed by the National Research Council [3], specifically recommended more attention in this area almost four decades ago. The MSBC as applied to two-and three-dimensional airfoils with flaps is the focus of the present study. 1

15 Figure 1-1 Examples of several devices leading to increased lift through delay in stall. Indicated values are arbitrary [1].

16 1.2 A Brief Review of the Relevant Literature Obviously, the forces and moments responsible for vibration of a wing are governed by the pressure distribution on its surface which, in turn, depend on the character of the boundarylayer and its separation. Hence, ever since the introduction of the boundary-layer concept by Prandtl (1904), there has been a constant challenge faced by scientists and engineers to minimize its adverse effects and control it to advantage. As pointed out earlier, several methods for control of the boundary-layer separation have been investigated at length and employed in practice with a measure of success. A vast body of literature accumulated over the years has been reviewed rather effectively by several authors including Goldstein [4], Lachmann [5], Rosenhead [6], Schlichting [7], Chang [8], and others. Irrespective of the method used, the main objective of a control procedure is to prevent, or at least delay, separation of the boundary-layer from the wall. A moving surface attempts to accomplish this in two ways: it prevents the initial growth of boundary-layer by minimizing relative motion between the surface and the free stream; and it injects momentum into the existing boundary-layer. The injection of momentum helps in keeping the flow attached to the surface in the region of adverse pressure gradient thus delaying the separation of the boundary-layer. Newton was probably the first one to observe the effect of moving wall boundary-layer control on the trajectory of a spinning ball [9], although the basis of the effect was not fully recognized. Almost 200 years later, Magnus [10] studied the lift generated by circulation and utilized the effect to construct a ship with a vertical rotating cylinder replacing the sail. Swanson [11] and Iverson [12] have presented excellent reviews of literature on the Magnus effect. As early as in 1910, Prandtl himself demonstrated his "ship of zero resistance" (Figure 1-2) through flow around two counter-rotating cylinders [13], while Flettner [14] applied the principle to ship 3

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18 propulsion in 1924 when he fitted large vertical rotating cylinders on the deck of the "Buchau" (Figure 1-3). A little later, in 1934, Goldstein [4] illustrated the principle of boundary-layer control using a rotating cylinder at the leading edge of a flat plate. However, the most practical application of a moving wall for boundary-layer control was demonstrated by Favre [15]. Using an airfoil with upper surface partly formed by a belt moving over two rollers (Figure 1-4), he was able to delay separation until the angle of attack reached 55 where the maximum lift coefficient of 3.5 was realized. After a lull of more than twenty years ( ), during which the tempo of research activity, as indicated by important contributions in the field, remained dormant, there appears to be some signs of renewed interest in this form of boundary-layer control. Alvarez-Calderon and Arnold [16] carried out tests on a rotating cylinder flap to evolve a high lift airfoil for STOL-type aircraft. The system was flight tested on a single engine high-wing research aircraft designed by Aeronautical Division of the Universidad Nacional de Ingenieria in Lima, Peru [17]. Around the same time, Brooks [18] presented his preliminary results of tests on a hydrofoil with a rotating cylinder at the leading or trailing edge. For the leading edge configuration only a small increase in lift was observed, however, for the latter case a substantial gain in lift resulted. Motivation for the test-program was to assess improvement in the fin performance for torpedo control. Along the same line, Steele and Harding [19] studied the application of rotating cylinders to improve ship-maneuverability. Extensive force measurements and flow visualization experiments were conducted using a water tunnel and a large circulating water channel. Three different configurations of rudder were used. The rotating Cylinder: (i) in isolation; 5

19 6

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21 (ii) (iii) at the leading edge of a rudder; combined with a flap-rudder, the cylinder being at the leading edge of the flap. From the overall consideration of hydrodynamic performance, mechanical complexity and power consumption, the configuration in (ii) was preferred. An application to a 250,000 ton tanker showed the power requirement for 1 m diameter cylinder rotating at 350 rpm to be around 400 kw. Of some interest is the North American Rockwell OV-10A (Figure 1-5) which was flighttested by NASA's Ames Research Center [20-22]. Cylinders, located at the leading edge of the flaps, are made to rotate at high speed with the flaps in lowered position. The main objective of the test-program was to assess handling qualities of the propeller-powered STOL-type aircraft at higher lift coefficients. The aircraft was flown at speeds of m/s, along approaches up to -8, which corresponded to the lift coefficient of about 4.3. In the pilot's opinion, any further reductions in the approach speed were limited by the lateral-directional stability and control characteristics. Excellent photographs of the airplane on the ground, showing the cylinders in position, and in flight have been published in the Aviation Week and Space Technology [23]. Around the same time Tennant presented an interesting analysis for the two dimensional moving wall diffuser with a step change in area [24,25]. The diffuser incorporated rotating cylinders to form a part of its wall at the station of the area change. Preliminary experiments were also conducted for the area ratio up to 1:2.5, which showed no separation for appropriate moving surface to diffuser inlet velocity ratio. Tennant et al. [26] have also conducted tests with a wedge shaped flap having a rotating cylinder as the leading edge. Flap deflection was limited to 15 and the critical cylinder velocity necessary to suppress separation was determined. Effects of increase 8

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23 in gap-size (between the cylinder and the flap surface) were also assessed. No effort was made to observe the influence of increase in cylinder surface velocity beyond Uc/U = 1.2. Subsequently, Tennant et al. [27] have reported circulation control for a symmetrical airfoil with a rotating cylinder forming its trailing edge. For zero angle of attack, the lift coefficient of 1.2 was attained with Uc/U = 3. Also of interest is their study concerning boundary-layer growth on moving surfaces accounting for gap effects [28,29]. It is important to point out that virtually all the studies used two-dimensional models in the investigations. Even the design of OV-10A was based on two-dimensional data! This was the state of development with respect to the moving surface boundary-layer control when Modi et al. entered the field in 1979 [30]. The subsequent contributions to the literature are essentially from his group which undertook planned, comprehensive investigations with two-dimensional airfoils (Figure 1-6) as well as two-and three-dimensional bluff bodies (Figure 1-7) involving a flat plate (two-dimensional), rectangular prism (two-dimensional), and tractor trailer truck configurations (three-dimensional). The extensive wind tunnel test-program, complemented by numerical simulations and flow visualization (Figure 1-8), showed: (i) significant increase in the lift of an airfoil by as much as 195%, with the stall delayed to 48 [31]!; (ii) reduction in the pressure drag of the flat plate by 75% [32]; (iii) reduction in the pressure drag of rectangular prisms by around 55% [32]; (iv) reduction in the pressure drag of highway truck models by 24% [33]. Modi has presented a detailed review of this class of problem [2]. It is of some interest to mention a recent contribution by Mokhtarian et al. [34] at Bombardier Aerospace. The study characterized the effects of the leading edge MSBC on the 10

24 The Basic Configuration Leading-Edge Cylinder Trailing-Edge Figure 1-6 Various rotating cylinder configurations studied with the twodimensional Joukowski airfoil model by Modi et al. [31]. 11

25 Two dimensional flat plate Two dimensional rectangular prisms a U H Uc - L L/H = 0.3,1,2,4 Tractor-trailer truck configuration U t Figure 1-7 Application of the moving surface boundary-layer control to bluff bodies such as a flat plate at large angles of attack, rectangular prisms, and tractor-trailer truck configurations [32,33]. 12

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27 development of the boundary-layer over a NACA 0015 airfoil. Measurements of the boundarylayer were made with a hot-wire probe, and both average as well as time-varying flow data were presented. The effect of momentum injection was primarily to increase the peak negative pressure. It delayed the boundary-layer flow separation, narrowed the wake, and increased the lift. It was also found that the momentum thickness of the airfoil boundary-layer not only decreased as cylinder rotation increased, but also grew linearly with the streamwise distance along the airfoil surface. 1.3 Scope of the Present Investigation The present investigation builds on this background. It presents the results of a wind tunnel test-program undertaken to assess effectiveness of the MSBC as applied to the high performance, low speed NASA LS(1)-0417 airfoil with a plain flap. It is important to emphasize that the choice of airfoil is of little consequence to illustrate improved performance in presence of the MSBC. Experiments are conducted to assess the performance of both the two-dimensional (i.e. infinite aspect ratio, 2-D), and of the three-dimensional (i.e. finite aspect ratio, 3-D) cases. The experimental study has four distinct phases. They characterize improved performance of an aircraft wing through the practical application of the MSBC as follows: a) two-dimensional airfoil with its leading edge formed by a circular cylindrical rotating element for momentum injection into the boundary-layer; b) the airfoil with a plain flap, leading edge of the flap formed by a momentum injection cylinder; c) the airfoil with a plain flap, leading edges of both (airfoil and the flap) formed by momentum injection cylinders; 14

28 d) extension to a finite aspect ratio wing with the same configurations as in (a), (b) and (c). The test-program may be considered quite comprehensive and realistic as it provides useful information for both infinite as well as finite aspect ratio wings. The results should serve as important base information for future studies in this general area. 15

29 2. MODEL AND TEST-PROGRAM 2.1 Introductory Remarks The wind tunnel test-program has two main objectives. Primarily, it attempts to assess the two-dimensional (2-D) and three-dimensional (3-D) fluid dynamical performance of an aircraft wing, without and with a plain flap, in presence of the Moving Surface Boundary-Layer Control. A single model was designed, compatible with the existing wind tunnel facilities in the Mechanical Engineering Department at the University of British Columbia. Besides efficient use of the resources, it helped provide consistent test-conditions. The model was made of Computer Numeric Controlled (CNC) milled aluminum sections that formed the airfoil-shaped surfaces. The model uses the NASA LS(1)-0417, high performance, low speed airfoil. The airfoil designation may be explained as follows: LS low speed; (1) first family; 04 design lift coefficient of 0.4; 17 % thickness at 1/3 chord. The model has a chord of m. The airfoil caries a plain flap with a chord equal to 1/3 of the airfoil, and the MSBC cylinders at the leading edge of the airfoil (wing cylinder) as well as at the leading edge of the flap (flap cylinder). The fluid dynamical parameters of interest are the pressure distribution, lift as well as drag forces, and stall angle of attack as affected by the wing and flap deflection (a, 8f, respectively) as well as momentum injection parameters U w /U and Uf/U. Here: U = free stream velocity; 16

30 U w = surface velocity of the wing cylinder; Uf = surface velocity of the flap cylinder. The rotating elements were powered by a pair of variable speed electric motors. The size of the rotating element at the leading edge of the wing is 38 mm diameter while that at the flap has 25 mm diameter. The cylinder dimensions were so selected as to match the local curvature of the airfoil and provide effective momentum injection. Schematic diagrams of the model are presented in Figure 2-1. The photograph in Figure 2-2 shows the model in the wind tunnel used for 2-D study. This wind tunnel was recently designated the "Parkinson Wind Tunnel" to honour its designer Professor GV. Parkinson. 2.2 Two-Dimensional Study During the 2-D study, the model was tested in a low speed, low turbulence return-type wind tunnel where the air speed can be varied from 0-35 m/s with a turbulence level less than 0.1%. A Betz rnicromanometer, with an accuracy of 0.2 mm of water, is used to measure the pressure differential across the contraction section of 7:1 ratio. The test-section velocity is calibrated against the above pressure differential. The rectangular cross-section, x m, is 2.6 m long and is provided with 45 comer fillets which vary from x m to x m to partly compensate for the boundary-layer growth. The spatial variation of the mean velocity in the test-section is less than 0.25%. The tunnel is powered by a 9 kw direct current motor driving a commercial axial flow fan with a Ward-Leonard system of speed control. Figure 2-3 shows an outline of the tunnel. 17

31 Figure 2-1 Schematic diagrams showing model configurations used during the test-program: (a) NASA LS(1)-0417 airfoil, camber, 17% thick at 1/3 chord; (b) leading edge of the wing formed by a rotating cylinder; (c) the airfoil with a plain flap, leading edge of the flap replaced with a rotating cylinder; (d) a rotating cylinder at the leading edge of the wing and the flap. 18

32 u c c i a c o sg i C CU a cu u a a cu c CN CU ' 4- o -C a, ca O i g 19

33 20

34 2.3 Approach to 3-D Study Three-dimensional experiments were carried out in an open circuit boundary-layer wind tunnel (BLWT) with a test-section of 2.44 x 1.6 x 24.4 m (Figure 2-4). The relatively large testsection permits the floor boundary-layer to attain a thickness as large as 0.6 m with appropriate distribution of surface roughness elements. In the present study, the model was mounted close to the entrance of the test-section where the turbulence intensity is lower than 0.4%. The BLWT has a contraction ratio of 4.7:1 and uses screens as well as a honeycomb to straighten and homogenize the flow as it accelerates towards the test-section. The tunnel has a stable speed range of m/s. The wind speed is measured with a Pitot static tube and an inclined manometer with an accuracy of 0.5 %. As pointed out before, the model used for the 3-D experiments is the same as that employed during the 2-D study, except for minor modifications near the tip to accommodate part of the drive system. A photograph of the model in the BLWT with flap deflection is shown in Figure 2-5. The model wing had a span of 0.88 m with an aspect ratio of 6, and the MSBC cylinders covered 90% of the span. Because of the large cross-section of the BLWT, the blockage was negligible, even at a = 90 ( 6.5 %). 2.4 Pressure and Force Measurements The model was subjected to both pressure distribution and direct force measurements. There were 42 pressure taps distributed over the circumference at the center span: 32 at the central pressure manifold on the wing, witii the remaining at the flap. The pressure taps are 0.5 mm holes in brass inserts aligned perpendicular to the airfoil surface. The pressure information at the leading edge of the airfoil, around the rotating cylinder, is quite important as it would significantly contribute to the lift. However, the measurements of 21

35 22

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37 pressures at the surface of the cylinder, rotating typically at 6,000-12,000 rpm, presents a challenge. The problem was resolved by keeping the pressure conveying tubes, located in narrow grooves, stationary while the cylinder rotated (Figure 2-6). Thus the pressure was measured not at the surface of the cylinder but in a close proximity of it. As the pressure taps remained immersed in the boundary-layer, this would provide a good estimate of the static pressure. The pressure was measured by a Scanivalve Pressure Transducing System, which connected to the taps on the model through 1.7 mm internal diameter polyethylene tubes. Signals from the Scanivalve were amplified and then recorded by a PC-based data acquisition system equipped with a 12-bit Data Translation Card (DT-2801) and specially developed Windows software. A schematic diagram of the experimental set-up is shown in Figure 2-7. The data acquisition software streamlined the test-procedure by automatically controlling the Scanivavle, compiling the data, and performing all the necessary data manipulations and calculations. Direct measurement of lift and drag forces was also carried out to help assess twodimensional character of the flow. These data were obtained using a custom designed x-y, airbearing supported, balance. Cylindrical air bushings by New-Way, with an internal diameter of 5.1 cm, were used to provide near-frictionless movements. Forces were monitored using Precision Transducers PT1000 load cells with a force limit of 7 kg and the resolution of around 0.05 N. This corresponds to the C L sensitivity of Signals from the PT1000 were amplified by an Iotech DBK-16 strain-gage card, and then recorded by an Iotech Daqboard 2000 analog to digital data acquisition card on a Windows-based PC. The rotating cylinders were made from aluminum, precision ground for a smooth surface finish, and balanced for high-speed rotation. A rotating element is supported by a pair of high speed bearings, housed in brackets at either end of the model, and connected to an externally 24

38 Figure 2-6 Arrangement for pressure measurements on the leading edge rotating cylinder. 25

39 Model Motor Variac Tachometer Test-Section Wind Scanivalve Control Rotary ^-i Scanivalve t o o C=3 P r e s s u r e Signal Conditioning Pressure Tap Lines S i 9 n a l Data Acquisition Figure 2-7 Experimental setup for the acquisition of pressure data. 26

40 located drive (1/4 hp, 3.8A, Variac controlled motor) by a Fenner Coupling. The cylinder speed was measured by a microprocessor-based tachometer system, accurate to 30 rpm. Figure 2-8 presents a photograph of the model, showing the motor drive, rotating cylinders, pressure tubes and Variac speed control. Tests were carried out over a range of wing and flap cylinder rotation speeds, corresponding to U w /U = 0, 1, 2, 3, 4 and Uf/U = 0, 1, 3. The tests included three flap deflection settings, 8f = 0, 20 and 45. Angle of attack typically ranged from -6 to well above the a stall. The Reynolds number during the test-program was held constant at ~

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42 3. TWO-DIMENSIONAL STUDY This chapter considers 2-D fluid dynamics of the NASA LS(1)-0417 airfoil as affected by the momentum injection. With two rotating elements, one at the leading edge of the wing (wing cylinder) and the other at the flap (flap cylinder), several variations are possible: (i) no momentum injection through either of the cylinders to obtain basic information; (ii) the MSBC applied only at the wing; (iii) the momentum injected at the flap only; (iv) the MSBC implemented at both the wing and the flap. In all the cases, the angle of attack varies from near zero lift to beyond oc sta u with step sizes in the range of 2-5. Three different flap deflection angles are considered: 8f = 0, 20 and 45. The momentum injection parameters are varied systematically in the range U w /U = 0-4 and U f /U = 0-3. As can be expected, with the variation of parameters, the amount of information obtained is quite extensive. What is important are the results useful in establishing trends and that is where the attention is directed. The relatively large angles of attack used in the experiments result in a considerable blockage of the wind tunnel test-section, from 21% at a = 30 to 30% at a = 45. The wall confinement leads to an increase in the local wind speed, at the location of the model, thus resulting in higher aerodynamic forces. Several approximate correction procedures have been reported in the literature to account for this effect. However, these approaches are mostly applicable to streamlined bodies with attached flows. A satisfactory procedure applicable to bluff bodies offering large blockage, and with separating shear layers, is still not available. In absence 29

43 of any reliable procedure to account for wall confinement effects in the present situation, the results are purposely presented in the uncorrected form. With rotation of the cylinder(s), the problem is further complicated. As shown by the pressure data and confirmed by flow visualization [32], the unsteady flow can be separating and reattaching over a large portion of the top surface. Thus the situation is rather challenging. Fortunately, the main objective of the study is to assess the effect of momentum injection, which has been observed to be essentially independent of the Reynolds number in the subcritical range [32]. Thus the mechanics of delay in the boundary-layer separation due to the MSBC is virtually unaffected by the blockage, which would raise the local Reynolds number. Furthermore, the main interest is in assessing relative influence of the momentum injection at a given blockage. 3.1 Reference Airfoil The first step is to obtain pressure distribution and force information for the NASA airfoil which can serve as reference to assess the influence of momentum injection and flap deflection. The reference airfoil is also referred to as "base airfoil" or "basic airfoil". Note, the basic airfoil (NASA LS(1)-0417) neither carries a flap nor the momentum injection at the airfoil nose. To capture these features as closely as possible, the NASA airfoil model was slightly modified as shown in Figure 3-1. As expected, there is no flap. The airfoil nose is replaced by a non-rotating circular cylinder with matching curvature. Furthermore, the gaps between the cylinders and the airfoil are sealed. Figure 3-2 presents the pressure distribution on the surface of the airfoil in terms of nondimensional parameters (Cp, pressure coefficient; distance along the chord, x/c). Due to the practical difficulty in locating pressure taps in the airfoil's cusp region, there is an apparent discontinuity in the plots near the trailing edge. Fortunately, this region contributes little to the 30

44 Gap Sealed G ap S e a l e d Figure 3-1 Schematic diagram of the model used to obtain reference information showing: (a) base airfoil; (b) modifications leading to equivalent airfoil. 31

45 H i 1 i i x/c Figure 3-2 Pressure distribution for the basic airfoil, NASA LS(1)-0417, showing stall around

46 forces and hence is of no consequence in the present discussion. Note, pressure information on both top and bottom surfaces of the airfoil are presented although it is the effect of momentum injection on the top surface that is of prime importance. For clarity, pressure plots on the lower surface of the airfoil are represented by dotted lines. This helps to distinguish them from the suction plots, on the top surface, which show the important outcome of the momentum injection. 3.2 MSBC Applied only at the Wing It is apparent from Figure 3-2 that, in absence of the MSBC, the NASA reference airfoil stalls at an angle of attack near 12. This is suggested by the collapse of the negative pressure peaks and the essentially flat pressure profile on the top surface. Force data presented later (Figure 3-6) also substantiate this observation. The bottom surface displays positive pressure near the nose as expected. A momentum injection changes the situation dramatically as shown in Figure 3-3. Note, now the angle of attack is 30. In absence of the MSBC, the airfoil should have stalled as seen in Figure 3-2. However, even with a relatively small amount of momentum injection of U w /U = 2, the flow clearly reattaches as indicated by the suction peak near the wing's leading edge. Further raising of the momentum injection to U w /U = 3 and 4, increases the magnitude of the suction peak and decreases the Cp near the trailing edge of the airfoil. Note, the boundary-layer separation is delayed to nearly 80% of the chord, and the airfoil displays essentially potential flow! It is remarkable that the boundary-layer separation moves downstream, even with the angle of attack as high as a = 40, to 70% of the chord, for U w /U = 4 (Figure 3-4). For the base airfoil only, lift and drag were determined through integration of the pressure distribution on the wing, as well as by direct measurement with the balance mentioned in Chapter 2. The results are compared in Figure 3-5. Here: C_, C D are lift and drag coefficients, 33

47 2 H 1 i 1 1 i 1 i x/c Figure 3-3 Pressure distribution for the airfoil in presence of the leading edge momentum injection. 34

48 x/c Figure 3-4 Pressure plots for the airfoil, with the momentum injection of U w /U = 4, at several angles of attack. 35

49 1

50 respectively, calculated using directly measured force values, while C L P and C D P are their pressure-based counterparts. It is clear from the figure that the lift coefficients as given by the pressure integration and the direct force measurements are in relatively good agreement. On the other hand, the drag coefficient in the range 0 < a < 12 is noticeably larger for the direct force measurement results. This conforms with the expected behavior. The pressure data do not account for the skin friction contribution. Of course, the direct force measurement gives the combination of pressure drag and the resistance induced by shear stress. The Cp data were integrated to obtain the lift and drag information as affected by the momentum injection parameter, U w /U, and angle of attack. Figure 3-6 summarizes the effects of wing cylinder rotation on C L. At the outset it is of interest to recognize that the momentum injection does not significantly change the average slope of the lift curve before stall. The base airfoil (i.e. the one with U w /U = 0) has a maximum lift coefficient of around 0.9 at the stall angle of 10. The stall sets in rather suddenly as seen earlier with the collapse of the negative pressure peak (Figure 3-2). However, with the cylinder rotation even at the lowest rate used (U w /U = 1), the lift and stall characteristics are significantly improved. The airfoil exhibits a desirable gradual onset of stall. The maximum lift coefficient obtained with U w /U = 4 is around 2.5 at a = 45, which is almost 2.8 times the C L of the base airfoil (increase in C_ m a x by 177%) with the stall delayed from 10 to 50. Only a marginal improvement in performance was obtained through an increase in U w /U beyond 4. This is understandable. Earlier, the pressure plots for U w /U = 4 showed the flow to be essentially potential, i.e. the boundary-layer separation is near the trailing edge, even at large angles of attack. So there is no significant scope for causing additional delay of the boundary- 37

51 38

52 layer separation by further injection of momentum. Additional injected momentum will be lost in the wake. Figure 3-7 presents typical drag data for the airfoil with momentum injection at the wing. Note, for a given U w /U, each curve shows a distinct increase in drag beyond the point of stall. These points are designated as 'S 0 ' for U W AJ = 0, "Si" for U W AJ = 1, etc. Figure 3-8 shows a close-up of the low angle of attack region covered by a in the range -6 to 20. It is apparent that the drag can approach a very low value, at small angles of attack, in presence of the momentum injection. Den Hertog in his study with underwater vehicles also came across a similar situation [35]. With the momentum injection of U w AJ = 3 at the leading edge of a NACA 0021 airfoil, the drag coefficient was virtually reduced to zero (C_ = 0.001)! 3.3 Rotating Cylinder at the Flap The model was next tested with the flap cylinder. Unlike the wing cylinder configuration, the flap with a rotating element changes the basic geometry substantially. The configuration may be referred to as a "rotating flap". Figure 3-9 shows pressure distribution for the airfoil with the MSBC applied at the flap. The results are for a = 6. It shows data for the base airfoil, the airfoil with 8f = 45 but no momentum injection, and 8f =45 with Uf AJ = 3. With this flap deflection, the airfoil stalled beyond a = 6. The figure clearly isolates the effects of flap deflection and momentum injection on the pressure distribution. For Uf AJ = Sf = 0, there is a relatively small suction peak (Cp = -1.8) which increases to -2.5 for 8f =45, and Uf AJ = 0; and attains a value of -3.8 with the addition of momentum 39

53 40

54 41

55 y i 0 'r v-a- 1 "_7 I x/c Figure 3-9 Pressure plots showing the effect of flap deflection and momentum injection. 42

56 injection. As expected, the effect of flap deflection is to shift the lift plots to the left (Figure 3-10). Note, the wing cylinder at the airfoil nose is not rotating. This helps to focus on the effect of Uf. Two cases are considered: 8f = 0 and 45, with Uf/U = 0, 1, 3. For the case of 8f = 0, the airfoil stalls at a ~ 10 as before, and the peak C L m a x is around With the flap deflection of 8f = 45, as anticipated the plots move to the left. The peak C L m a x now is around 2.25 with the stall setting in at a = 6. It is of interest to recognize that at a = 3, the increase in C_ m a x for Uf/U = 3 is 300%. This is indeed remarkable. It can be used to advantage during landing maneuvers. 3.4 Momentum Injection at the Wing and the Flap The next logical step was to assess the effect of rotation of both the cylinders. The pressure distribution as affected by momentum injection at the wing, with a = 5, Uf/U = 3 and 8f = 45 is shown in Figure The figure also shows data for a = 0 as reference. Here the effect of U w is to shift the negative peak Cp towards the leading edge of the wing and to increase the suction pressure over the entire topi surface. Figure 3-12 shows the effect of momentum injection at the flap on the pressure distribution for the airfoil at an angle of attack of 30 with U w /U = 4 and 8f = 45. Results for the airfoil with U w /U = 4 and 8f = 0 are included for comparison. The flap cylinder rotation causes the peak suction Cp to increase from -8.8 to nearly -10 at the leading edge of the wing. It also increases the suction pressure over the entire surface of the wing, especially near the leading edge of the flap. The variation in C_ with a is presented in Figure Three values of the flap deflection are considered, 8f = 0, 20 and 45, with Uf/U = 0 and 3, and U w /U = 4. As before, the base airfoil data are included to help with comparison. The peak C L m a x is around 3.6, an increase of approximately 300% over that for the base airfoil, and a rise of about 50% over the performance 43

57 I i -1H i i i i i n a Figure 3-10 Effect of flap deflection and cylinder rotation on the variation of C L with a.

58 Re = 1fr Uf/U = 3, 8 f = 45 u w /u = o, a = 0 u w /u = o, a = 5 -o- u w /u = 4, a = 5 x/c Figure 3-11 Effect of wing cylinder rotation on the pressure distribution for the airfoil with U f /U = 3 and 8 f =

59 -10 Re = 10 5 Uw/U = 4 a = 30 Uf/U = o, 5f = Uf/U = o,5f = 45 Uf/U = 3,8f = 45 o x/c Figure 3-12 Effect of momentum injection at the flap on the pressure distribution for the airfoil with U w AJ = 4 and 8 f = 45. The pressure distribution for U w AJ = 4 and 8f = 0 is included for reference. 46

60 a 0 ure 3-13 Variation of C L with a for the airfoil with both rotating cylinders and flap deflection. 47

61 of the airfoil with wing cylinder rotation alone (Figure 3-13). The stall is delayed from a = 10 to 35 for 8f = 45, Uf/U = 3 and U w /U = 4. The enhanced lift due to flap cylinder rotation is clearly displayed, e.g. the airfoil with 8f = 20 and Uf/U = 3. Again, the effect of the flap deflection and flap cylinder rotation is to shift the lift curve to the left, with a decrease in the a stall from 45 to 35. This is in contrast to the effect of leading-edge cylinder rotation. As seen before (Figure 3-6), it extends the lift curve by delaying separation of the boundary-layer over the top surface of the airfoil. Figure 3-14 presents typical drag data in presence of both the cylinder rotation. Five cases are considered: 8f = 0, 20, 45 with Uf/U = 0, 3. U w /U is held fixed at 4. Again, the base airfoil data are included for comparison. The region of particular interest extends from a = -10 to 20. As observed earlier, the C D again becomes extremely small, at low angles of attack, suggesting significant drag reduction. Furthermore, the drag penalty for the case 8f = 20 and Uf/U = 3 is slightly less than that for the case of 8f = 45 and Uf/U = 0, even though, as observed earlier (Figure 3-13), the lift is slightly higher. The two rotating cylinders combine their individual contributions for a remarkable effect on the airfoil performance. The leading edge cylinder delays separation of the boundary-layer, allowing the airfoil to operate at high angles of incidence without stall. Momentum injection at the deflecting flap was found to increase the suction peak at both higher 8f and Uf/U (Figure 3-9). This would then reflect on higher lift (Figure 3-10). Furthermore, the lift plots move to the left and hence the effects are felt at angles of attack which are smaller by the amount of the shift. It is apparent that the combination of the MSBC applied at the wing and flap lead to an airfoil with an outstanding lift and drag performance. Appendix A summarizes some of these results. 48

62 Basic Airfoil - Uf/U = 0, 8f = 0 Re = 10 5 U w /U = 4 - -m - - Uf/U = 0, 8f = D - - Uf/U = 3, 5f = 20 _ Uf/U = 0, 8f = 45 - Uf/U = 3, 5f = Figure 3-14 Typical drag data for the airfoil in the presence of both rotating cylinders andflapdeflection. 49

63 4. THREE-DIMENSIONAL STUDY The purpose of the 3-D study was to discover, in general, how the behavior of the airfoil with flap and the MSBC is affected by a finite aspect ratio. Of course, the primary effect of the 3-D character is the wing-tip vortex, which is due to the pressure differential between the bottom and top wing surfaces at the tip. The main objective is to understand basic trends in moving from the 2-D to 3-D study. Both the pressure distribution at center span as well as the variation of lift and drag with angle of attack are compared with the 2-D case to establish 3-D effects. Again, the angle of attack is varied from near zero-lift to beyond a st _j in step sizes between 2 and 5. For the 3-D investigation, force components are determined from direct measurements with the air bearingbased x-y balance described in Chapter 2. The aspect ratio was fixed at six, and the Reynolds number at 10 5, which matches the Re of the 2-D study. The flap and rotating elements covered 90% of the span from the root. The remaining 10% near the wing tip used the unmodified NASA LS(1)-0417 airfoil. It acted as a fairing for the bearing bracket and wing-support structure. 4.1 Base NASA LS(1) Airfoil Figure 4-1 shows the variation of pressure distribution at the center span with angle of attack for the 3-D base airfoil. Note, the base airfoil implies the original NASA profile without flap and momentum injection. There is the usual negative pressure peak near the leading edge that grows stronger as a increases. At a ~ 12, the pressure on the suction surface collapses and the wing is stalled. The trend is similar to that observed for the 2-D case, but there is a difference in the magnitude of Cp on the suction surface of the wing. This becomes apparent at angles of attack 50